Many airframe structural assemblies such as an aircraft fuselage are commonly-fabricated using a technique known as semi-monocoque construction. In this construction technique, high strength materials such as aluminum alloys are used to form an outer skin. Various substructure elements such as stiffeners are attached to the outer skin. The outer skin is typically designed to carry tension loads while the stiffeners are typically designed to carry bending and compression loads. The combination of the outer skin and the attached stiffeners results in a high-strength, low-weight structure that is well-suited for use in aircraft and other vehicles.
Fiber metal laminates are increasingly employed in semi-monocoque construction as a substitute for monolithic or homogeneous metallic skins. Typically comprised of alternating layers of metallic plies bonded to composite or fiberglass plies, fiber metal laminates are typically lighter in weight than homogeneous metallic skins of comparable thickness. In addition, for comparable skin thicknesses, fiber metal laminates exhibit improved mechanical properties such as increased fatigue life relative to homogeneous metallic skins.
Despite the advantages provided by fiber metal laminate skins, it is typically necessary to employ stringers and/or stiffeners with the skin as a means for carrying the above-mentioned bending and compression loads. A common method for attaching stiffeners to skins is with mechanical fasteners such as rivets, hi-locks and various other fastening systems. In this regard, it is typically necessary to drill a large quantity of fastener holes through the skin and stiffeners in order to allow for installation of the mechanical fasteners.
For large aircraft, thousands of such fastener holes must typically be drilled and an equal number of fasteners must be installed in a time-consuming and labor-intensive process in order to attach the stiffeners to the skin. Although mechanical fastening of stiffeners to skins has been a satisfactory method for its intended purpose, such method presents certain drawbacks which, over time, can increase operating and maintenance costs and can reduce service life.
For example, it is well-known in the art that holes in a load-carrying member are typically locations of increased stress in the member. Particularly in primary load-carrying members of an aircraft such as the aircraft fuselage, skins are subjected to repeated applications of high operating loads which, in turn, results in the repeated concentration of localized stresses at the fastener holes. Over time, such localized stress concentrations may lead to the occurrence of undesirable effects in the skin. Although fiber metal laminate skins tend to limit the manifestation of such undesirable effects as compared to homogeneous metallic skins, such undesirable effects may nonetheless occur.
As can be seen, there exists a need in the art for a system and method for attaching substructure elements such as stiffeners to fiber metal laminate skins without the use of mechanical fasteners. Furthermore, there exists a need in the art for a system and method for joining substructure elements to a fiber metal laminate skin with reduced fabrication and assembly costs and in a reduced amount of time such that the extended fatigue life capabilities of fiber metal laminates can be employed to their full benefit.
The features, functions and advantages that have been discussed can be achieved independently in various embodiments of the present disclosure or may be combined in yet other embodiments, further details of which can be seen with reference to the following description and drawings below.